Composite fan blade

ABSTRACT

A blade fabrication method is provided and includes additively manufacturing a core, securing the core to a mandrel, electroforming a leading edge sheath directly onto the core and the mandrel and removing the mandrel from the core and the leading edge sheath.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a Divisional of application Ser. No. 16/209,448filed, Dec. 4, 2018, the disclosure of which is incorporated herein byreference in its entirety.

BACKGROUND

Exemplary embodiments of the present disclosure relate generally tocomposite fan blades and, in one embodiment, to a composite fan bladewith a fabricated leading edge sheath.

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate ahigh-energy exhaust gas flow. The high-energy exhaust gas flow expandsthrough the turbine section to drive the compressor and the fan section.The compressor section typically includes low and high pressurecompressors and the turbine section includes low and high pressureturbines.

Within the compressor section, high energy fluids aerodynamicallyinteract with blades and vanes such that air flowing into the gasturbine engine can be compressed. Likewise, within the turbine section,high energy fluids, such as the products of combustion, aerodynamicallyinteract with blades and vanes in order to expand and to thereby drivecompressor and rotor rotation.

The blades in the turbine section in particular are typically exposed tohigh-temperatures and pressures and need to be structurally sound andcooled. As such, composite fan blades have been proposed and designed toserve as blades for turbine sections of gas turbine engines. Suchcomposite fan blades can include a leading edge sheath to meet designtargets. The leading edge sheath can be formed using leading edgedeposition processes or direct laser metal sintering (DLMS) processesbut it has been found that each of these has certain limitations.

BRIEF DESCRIPTION

According to an aspect of the disclosure, a blade fabrication method isprovided and includes additively manufacturing a core, securing the coreto a mandrel, electroforming a leading edge sheath directly onto thecore and the mandrel and removing the mandrel from the core and theleading edge sheath.

In accordance with additional or alternative embodiments, the additivelymanufacturing comprises direct metal laser sintering (DMLS).

In accordance with additional or alternative embodiments, the additivelymanufacturing includes additively manufacturing the core to be one ormore of solid, perforated and micro-latticed.

In accordance with additional or alternative embodiments, the securingincludes inserting an alignment pin into the core and the mandrel.

In accordance with additional or alternative embodiments, theelectroforming includes electroforming the leading edge sheath toinclude an elongate leading edge portion that extends forwardly from aleading edge of the core and sidewall portions that extend rearwardlyfrom a trailing edge of the elongate leading edge portion along the coreand a forward portion of the mandrel.

In accordance with additional or alternative embodiments, the methodfurther includes locally thickening the leading edge sheath tofacilitate retention of the core by the leading edge sheath.

In accordance with additional or alternative embodiments, the removingincludes inserting one or more wedges between the leading edge sheathand the mandrel.

In accordance with additional or alternative embodiments, the methodfurther includes bonding a blade body to the core and the leading edgesheath and the bonding includes adhering the blade body to the core andthe leading edge sheath.

According to another aspect of the disclosure, a method of fabricating ablade for use in a flowpath is provided and includes additivelymanufacturing a core having a length sufficient to span a substantialfraction of the flowpath, securing the core to a mandrel having a lengthsufficient to span the substantial fraction of the flowpath,electroforming a leading edge sheath directly onto the core and themandrel, separating the mandrel from the leading edge sheath along anentirety of the length of the mandrel and removing the mandrel from thecore and the leading edge sheath.

In accordance with additional or alternative embodiments, the additivelymanufacturing includes direct metal laser sintering (DMLS).

In accordance with additional or alternative embodiments, the additivelymanufacturing includes additively manufacturing the core to be one ormore of solid, perforated and micro-latticed.

In accordance with additional or alternative embodiments, the securingincludes inserting an alignment pin into the core and the mandrel.

In accordance with additional or alternative embodiments, theelectroforming includes electroforming the leading edge sheath toinclude an elongate leading edge portion that extends forwardly from aleading edge of the core and sidewall portions that extend rearwardlyfrom a trailing edge of the elongate leading edge portion along the coreand a forward portion of the mandrel.

In accordance with additional or alternative embodiments, the methodfurther includes locally thickening the leading edge sheath tofacilitate retention of the core by the leading edge sheath.

In accordance with additional or alternative embodiments, the removingincludes inserting one or more wedges between the leading edge sheathand the mandrel.

In accordance with additional or alternative embodiments, the methodfurther includes bonding a blade body to the core and the leading edgesheath and the bonding includes adhering the blade body to the core andthe leading edge sheath.

According to another aspect of the disclosure, a leading edge sheathassembly for a blade is provided and includes an additively manufacturedcore and a leading edge sheath. The additively manufactured coreincludes a leading edge, a trailing edge and first and second sidewallsextending from opposite sides of the leading edge to opposite sides ofthe trailing edge. The leading edge sheath is electroformed directlyonto the core and includes an elongate leading edge portion that extendsforwardly from the leading edge of the core and sidewall portions thatextend rearwardly from a trailing edge of the elongate leading edgeportion along and beyond the first and second sidewalls of the core.

In accordance with additional or alternative embodiments, the core isone or more of solid, perforated and micro-latticed.

In accordance with additional or alternative embodiments, a fan blade isprovided and includes the leading edge sheath assembly, a blade body andadhesive disposed to adhere interior surfaces of the sidewall portionsto sidewalls of the blade body and the trailing edge of the core to aleading edge of the blade body.

These and other advantages and features will become more apparent fromthe following description taken in conjunction with the drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The following descriptions should not be considered limiting in any way.With reference to the accompanying drawings, like elements are numberedalike:

FIG. 1 is a partial cross-sectional view of a gas turbine engine;

FIG. 2 is a partial cross-sectional view of an embodiment of a portionof a compressor section of the gas turbine engine of FIG. 1;

FIG. 3 is a partial cross-sectional view of another embodiment of aportion of a compressor section of the gas turbine engine of FIG. 1;

FIG. 4 is a flow diagram illustrating a method of fabricating acomposite fan blade in accordance with embodiments;

FIG. 5 is a perspective view of a composite fan blade in accordance withembodiments;

FIG. 6 is a diagram illustrating a method of fabricating a leading edgesheath of a composite fan blade in accordance with embodiments;

FIG. 7 is a cross-sectional view of a composite fan blade with a leadingedge sheath in accordance with embodiments;

FIG. 8 is a cross-sectional view of a composite fan blade with a leadingedge sheath in accordance with embodiments;

FIG. 9 is an enlarged cross-sectional view of a core of a leading edgesheath of a composite fan blade in accordance with embodiments;

FIG. 10 is an enlarged cross-sectional view of a locally thickenedportion of a leading edge sheath for core retention in accordance withembodiments; and

FIG. 11 is a cross-sectional view illustrating a separation of sidewallportions of a leading edge sheath of a composite fan blade from amandrel using wedges in accordance with embodiments.

These and other advantages and features will become more apparent fromthe following description taken in conjunction with the drawings.

DETAILED DESCRIPTION

A detailed description of one or more embodiments of the disclosedapparatus and method are presented herein by way of exemplification andnot limitation with reference to the Figures.

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude other systems or features. The fan section 22 drives air along abypass flow path B in a bypass duct, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 and then expansion through the turbinesection 28. Although depicted as a two-spool turbofan gas turbine enginein the disclosed non-limiting embodiment, it should be understood thatthe concepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary gas turbine engine 20 generally includes a low speed spool30 and a high speed spool 32 mounted for rotation about an enginecentral longitudinal axis A relative to an engine static structure 36via several bearing systems 38. It should be understood that variousbearing systems 38 at various locations may alternatively oradditionally be provided, and the location of bearing systems 38 may bevaried as appropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in the gas turbineengine 20 between the high pressure compressor 52 and the high pressureturbine 54. The engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. The enginestatic structure 36 further supports the bearing systems 38 in theturbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 andthen the high pressure compressor 52, is mixed and burned with fuel inthe combustor 56 and is then expanded over the high pressure turbine 54and the low pressure turbine 46. The high and low pressure turbines 54and 46 rotationally drive the low speed spool 30 and the high speedspool 32, respectively, in response to the expansion. It will beappreciated that each of the positions of the fan section 22, compressorsection 24, combustor section 26, turbine section 28, and fan drive gearsystem 48 may be varied. For example, geared architecture 48 may belocated aft of the combustor section 26 or even aft of the turbinesection 28, and the fan section 22 may be positioned forward or aft ofthe location of geared architecture 48.

The gas turbine engine 20 in one example is a high-bypass gearedaircraft engine. In a further example, the gas turbine engine 20 bypassratio is greater than about six (6), with an example embodiment beinggreater than about ten (10), the geared architecture 48 is an epicyclicgear train, such as a planetary gear system or other gear system, with agear reduction ratio of greater than about 2.3 and the low pressureturbine 46 has a pressure ratio that is greater than about five. In onedisclosed embodiment, the gas turbine engine 20 bypass ratio is greaterthan about ten (10:1), the fan diameter is significantly larger thanthat of the low pressure compressor 44, and the low pressure turbine 46has a pressure ratio that is greater than about five 5:1. Low pressureturbine 46 pressure ratio is pressure measured prior to inlet of lowpressure turbine 46 as related to the pressure at the outlet of the lowpressure turbine 46 prior to an exhaust nozzle. The geared architecture48 may be an epicycle gear train, such as a planetary gear system orother gear system, with a gear reduction ratio of greater than about2.3:1. It should be understood, however, that the above parameters areonly exemplary of one embodiment of a geared architecture engine andthat the present disclosure is applicable to other gas turbine enginesincluding direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the gas turbine engine 20is designed for a particular flight condition—typically cruise at about0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).

Referring now to FIG. 2, either or both of the low pressure compressor44 or the high pressure compressor 52 includes a compressor case 60, inwhich compressor rotors 62 are arranged along an engine axis 64 aboutwhich the compressor rotors 62 rotate. Each compressor rotor 62 includesa rotor disc 66 with a platform 70 and a plurality of rotor blades 68extending radially outwardly from the platform 70 (i.e., a rotor stack).In some embodiments, the rotor disc 66 and the plurality of rotor blades68 are a single, unitary structure, an integrally bladed compressorrotor 62. In other embodiments, the rotor blades 68 are each installedto the rotor disc 66 via, for example, a dovetail joint where a tab orprotrusion at the rotor blade 68 is inserted into a corresponding slotin the platform 70.

As shown in FIG. 2, axially adjacent compressor rotors 62 may be joinedto each other, while in other embodiments, as shown in FIG. 3, thecompressor rotor 62 may be joined to another rotating component, such asa spacer 72. The compressor rotor 62 is secured to the adjacent rotatingcomponent by an interference fit or a “snap fit,” which in someembodiments is combined with another mechanical fastening, such as aplurality of bolts (not shown) to secure the components and to form ordefine a snap location.

As will be described below, a composite fan blade is provided for use ina flowpath, such as a flowpath of one or more of the fan section 22, thecompressor section 24, the combustor section 26 and the turbine section28 of the gas turbine engine 20 described above. The composite fan bladeincludes a leading edge sheath that is formed by an improved leadingedge deposition process combined with a direct metal laser sinterin(DMLS) process. The composite fan blade can thus be lightweight. Thecomposite fan blade has a forward nose that is filled with a solid orsemi-solid (e.g., perforated or micro-latticed) core, which isadditively manufactured or formed from the DMLS process, and a leadingedge sheath or outer skin that is formed from a uniform or a slightlynon-uniform deposition process. The core serves to address straincapabilities of leading edge of the composite fan blade. In addition,the presence of the core allows for the leading edge sheath or outerskin to be deposited directly within the capabilities of the depositionprocess.

With reference to FIGS. 4, 5 and 6, a method of fabricating a compositefan blade 501 is provided such that the composite fan blade 501 can beinstalled in a flowpath, such as a flowpath of one or more of the fansection 22, the compressor section 24, the combustor section 26 and theturbine section 28 of the gas turbine engine 20 described above. Themethod includes additively manufacturing a core 610 (see FIG. 6) to havea length that extends along the length-wise dimension LD and issufficient to span a substantial fraction of the flowpath into which thecomposite fan blade 501 is to be installed (401). The method furtherincludes securing the core 610 to a mandrel 620 (see FIG. 6) where themandrel 620 has a length that also extends along the length-wisedimension LD and which is also sufficient to span the substantialfraction of the flowpath (402). In addition the method includeselectroforming a leading edge sheath 630 (see FIG. 6) directly onto thecore 610 and the mandrel 620 (403), separating the mandrel 620 from theleading edge sheath 630 along an entirety of the length of the mandrel620 (404) and removing the mandrel 620 from the core 610 and the leadingedge sheath 630 (405).

In accordance with embodiments and with reference to FIGS. 7, 8 and 9,the additive manufacturing of the core 610 of operation 401 can includeDMLS processing or another similar type of additive manufacturing andcan result in the core 610 being one or more of solid 611 (see FIGS. 7and 9), perforated 612 (see FIG. 9) and micro-latticed 613 (see FIGS. 8and 9). That is, as shown in FIG. 9, a first portion of the core 610(e.g., a central portion 901) can be solid 611, a second portion of thecore 610 (e.g., an outer portion 902) can be perforated 612 and a thirdportion of the core 610 (e.g., an intermediate portion 903 between thecentral portion 901 and the outer portion 902) can be micro-latticed613.

With reference back to FIG. 6, the core 610 is additively manufacturedto have a core body 614 with the above-noted length extending along thelength-wise dimension LD (see FIG. 5) and with a leading edge 615, atrailing edge 616 opposite the leading edge 615, a pressure surface orfirst side 617 and a suction surface or second side 618. The first andsecond sides 617 and 618 extend from opposite sides of the leading edge615 to opposite sides of the trailing edge 616. The core body 614 canhave a taper from the trailing edge 616 toward the leading edge 615 tofacilitate the formation of a leading edge sheath assembly having anose-cone shape.

The mandrel 620 is provided with a mandrel body 621 that can mimic theoverall shape 502 of the composite fan blade 501. The mandrel body 621has the above-noted length extending along the length-wise dimension LD(see FIG. 5) and a leading edge 622, a pressure surface or first side623 and a suction surface or second side 624. The first and second sides623 and 624 extend rearwardly from opposite sides of the leading edge622. The mandrel body 621 can have a taper toward the leading edge 622to facilitate the formation of a leading edge sheath assembly having across-sectional nose-cone shape.

With continued reference to FIG. 6 and with reference back to FIG. 4,the securing of the core 610 to the mandrel 620 of operation 402 can beexecuted such that the trailing edge 616 of the core 610 is secured tothe leading edge 622 of the mandrel 620. In addition, the securing ofoperation 402 can be achieved by inserting an alignment pin 640 into thetrailing edge 616 of the core 610 and the leading edge 622 of themandrel 620. It is to be understood that the alignment pin 640 can beprovided as multiple alignment pins 640 and that the multiple alignmentpins 640 can each be inserted into the trailing edge 616 of the core 610and the leading edge 622 of the mandrel 620 along the respective lengthsthereof.

With continued reference to FIG. 6 and with reference back to FIG. 4,the electroforming of the leading edge sheath 630 of operation 403 canbe executed directly on the core 610 and the mandrel 620 to retainfront/forward solidity for in-service reparability. The leading edgesheath 630 thus includes an elongate leading edge portion 631 andsidewall portions 632. The elongate leading edge portion 631 extendsforwardly from the leading edge 615 of the core 610. The sidewallportions 632 extend rearwardly from a trailing edge 633 of the elongateleading edge portion 631 (the trailing edge 633 can be defined as beingcoplanar with the leading edge 615 of the core 610), along the first andsecond sides 617 and 618 of the core 610 and along at least respectiveforward portions of the first and second sides 623 and 624 of themandrel 620. Exterior surfaces of the elongate leading edge portion 631and the sidewall portions 632 are smooth and continuous and taperedsimilarly to the tapers of the core 610 and the mandrel 620 such thatthe leading edge sheath 630 has the above-noted cross-sectionalnose-cone shape.

With continued reference to FIGS. 4 and 6 and with additional referenceto FIG. 10, the electroforming of the leading edge sheath 630 ofoperation 403 can include locally thickening the leading edge sheath 630to facilitate retention of the core 610 by the leading edge sheath 630(4031). As shown in FIG. 10, such local thickening can result in thesidewall portions 632 of the leading edge sheath 630 including athickened section 1001 which extends partially around the trailing edge616 of the core 610. This thickened section 1001 can serve to retain thecore 610 within the leading edge sheath 630 when the mandrel 620 isseparated and then removed from the leading edge sheath 630 inoperations 404 and 405.

With continued reference to FIGS. 4 and 6 and with additional referenceto FIG. 11, the separating of the mandrel 620 from the leading edgesheath 630 along an entirety of the length of the mandrel 620 ofoperation 404 can include an insertion of one or more wedges 1101between the sidewall portions 632 of the leading edge sheath 630 and thefirst and second sides 623 and 624 of the mandrel 620.

With reference back to FIG. 4 and to FIGS. 7 and 8, once the mandrel 620is removed from the core 610 and the leading edge sheath 630 inoperation 405, a blade body 650 (see FIGS. 7 and 8) can be bonded to thetrailing edge 616 of the core 610 and to interior surfaces of thesidewall portions 632 of the leading edge sheath 630 (406). The bondingcan be provided by an adhesive 651, such as a suitable epoxy or anothersimilar bonding adhesive.

As shown in FIGS. 7, 8 and 9, as a result of the method described above,a leading edge sheath assembly 701 can be provided for a blade, such asthe composite fan blade 501 of FIG. 5. The leading edge sheath assembly701 includes the additively manufactured core 610 and the leading edgesheath 630. The core 610 is one or more of solid 611 (see FIGS. 7 and9), perforated 612 (see FIG. 9) and micro-latticed 613 (see FIGS. 8 and9) and includes the leading edge 615, the trailing edge 616 and thefirst and second sidewalls 617 and 618. The leading edge sheath 630 iselectroformed directly onto the core 610 and includes the elongateleading edge portion 631 and the sidewall portions 632.

In accordance with embodiments, the leading edge sheath assembly 701 orthe blade as a whole (i.e., the composite fan blade 501 of FIG. 5) canalso include the blade body 650 and the adhesive 651.

Benefits of the features described herein are the provision of alightweight composite fan blade 501 through easy and tailorablefabrication methods and processes.

The term “about” is intended to include the degree of error associatedwith measurement of the particular quantity based upon the equipmentavailable at the time of filing the application.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the presentdisclosure. As used herein, the singular forms “a”, “an” and “the” areintended to include the plural forms as well, unless the context clearlyindicates otherwise. It will be further understood that the terms“comprises” and/or “comprising,” when used in this specification,specify the presence of stated features, integers, steps, operations,elements, and/or components, but do not preclude the presence oraddition of one or more other features, integers, steps, operations,element components, and/or groups thereof.

While the present disclosure has been described with reference to anexemplary embodiment or embodiments, it will be understood by thoseskilled in the art that various changes may be made and equivalents maybe substituted for elements thereof without departing from the scope ofthe present disclosure. In addition, many modifications may be made toadapt a particular situation or material to the teachings of the presentdisclosure without departing from the essential scope thereof.Therefore, it is intended that the present disclosure not be limited tothe particular embodiment disclosed as the best mode contemplated forcarrying out this present disclosure, but that the present disclosurewill include all embodiments falling within the scope of the claims.

What is claimed is:
 1. A blade fabrication method, comprising:additively manufacturing a core; securing the core to a mandrel;electroforming a leading edge sheath directly onto the core and themandrel; and removing the mandrel from the core and the leading edgesheath.
 2. The blade fabrication method according to claim 1, whereinthe additively manufacturing comprises direct metal laser sintering(DMLS).
 3. The blade fabrication method according to claim 1, whereinthe additively manufacturing comprises additively manufacturing the coreto be latticed in interior portions thereof and perforated at outerportions thereof.
 4. The blade fabrication method according to claim 1,wherein the securing comprises inserting an alignment pin into the coreand the mandrel.
 5. The blade fabrication method according to claim 1,wherein the electroforming comprises electroforming the leading edgesheath to comprise: an elongate leading edge portion that extendsforwardly from a leading edge of the core; and sidewall portions thatextend rearwardly from a trailing edge of the elongate leading edgeportion along the core and a forward portion of the mandrel.
 6. Theblade fabrication method according to claim 5, further comprisinglocally thickening the sidewall portions of the leading edge sheatharound a trailing edge of the core to facilitate retention of the coreby the leading edge sheath.
 7. The blade fabrication method according toclaim 1, wherein the removing comprises inserting one or more wedgesbetween the leading edge sheath and the mandrel.
 8. The bladefabrication method according to claim 1, further comprising bonding ablade body to the core and the leading edge sheath, wherein the bondingcomprises adhering the blade body to the core and the leading edgesheath.
 9. A method of fabricating a blade for use in a flowpath,comprising: additively manufacturing a core having a length sufficientto span a substantial fraction of the flowpath; securing the core to amandrel having a length sufficient to span the substantial fraction ofthe flowpath; electroforming a leading edge sheath directly onto thecore and the mandrel; separating the mandrel from the leading edgesheath along an entirety of the length of the mandrel; and removing themandrel from the core and the leading edge sheath.
 10. The methodaccording to claim 9, wherein the additively manufacturing comprisesdirect metal laser sintering (DMLS).
 11. The method according to claim9, wherein the additively manufacturing comprises additivelymanufacturing the core to be latticed in interior portions thereof andperforated at outer portions thereof.
 12. The method according to claim9, wherein the securing comprises inserting an alignment pin into thecore and the mandrel.
 13. The method according to claim 9, wherein theelectroforming comprises electroforming the leading edge sheath tocomprise: an elongate leading edge portion that extends forwardly from aleading edge of the core; and sidewall portions that extend rearwardlyfrom a trailing edge of the elongate leading edge portion along the coreand a forward portion of the mandrel.
 14. The method according to claim13, further comprising locally thickening the sidewalls portions of theleading edge sheath around a trailing ede of the core to facilitateretention of the core by the leading edge sheath.
 15. The methodaccording to claim 9, wherein the removing comprises inserting one ormore wedges between the leading edge sheath and the mandrel.
 16. Themethod according to claim 9, further comprising bonding a blade body tothe core and the leading edge sheath, wherein the bonding comprisesadhering the blade body to the core and the leading edge sheath.
 17. Afan blade formed by the method according to claim
 16. 18. A method ofassembling a leading edge sheath assembly for a blade having a bladebody, the method comprising: additively manufacturing a core perforatedat outer portions thereof and latticed in interior portions thereof, thecore comprising leading and trailing edges and first and secondsidewalls extending between the leading and training edges; securing thecore to a mandrel; electroforming a leading edge sheath directly ontothe core and the mandrel, the leading edge sheath comprising an elongateleading edge portion extending forwardly from the leading edge of thecore and sidewall portions extending rearwardly from the elongateleading edge portion along and beyond the first and second sidewalls ofthe core and comprising a thickened section extending partially aroundthe trailing edge of the core; separating the mandrel from the leadingedge sheath; and removing the mandrel from the core.
 19. The methodaccording to claim 18, wherein the separating of the mandrel from theleading edge sheath comprises inserting one or more wedges between theleading edge sheath and the mandrel.
 20. The method according to claim18, further comprising bonding a blade body to the core and the leadingedge sheath, wherein the bonding comprises adhering the blade body tothe core and the leading edge sheath.